EP3569823B1 - Seal assembly with baffle for gas turbine engine - Google Patents
Seal assembly with baffle for gas turbine engine Download PDFInfo
- Publication number
- EP3569823B1 EP3569823B1 EP19175130.4A EP19175130A EP3569823B1 EP 3569823 B1 EP3569823 B1 EP 3569823B1 EP 19175130 A EP19175130 A EP 19175130A EP 3569823 B1 EP3569823 B1 EP 3569823B1
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- EP
- European Patent Office
- Prior art keywords
- section
- seal
- region
- baffle
- cooling
- Prior art date
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Images
Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
- F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
- F01D11/005—Sealing means between non relatively rotating elements
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
- F01D25/08—Cooling; Heating; Heat-insulation
- F01D25/12—Cooling
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
- F01D25/24—Casings; Casing parts, e.g. diaphragms, casing fastenings
- F01D25/246—Fastening of diaphragms or stator-rings
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
- F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
- F01D11/14—Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing
- F01D11/20—Actively adjusting tip-clearance
- F01D11/24—Actively adjusting tip-clearance by selectively cooling-heating stator or rotor components
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2220/00—Application
- F05D2220/30—Application in turbines
- F05D2220/32—Application in turbines in gas turbines
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/10—Stators
- F05D2240/11—Shroud seal segments
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/10—Stators
- F05D2240/12—Fluid guiding means, e.g. vanes
- F05D2240/126—Baffles or ribs
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/55—Seals
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2250/00—Geometry
- F05D2250/10—Two-dimensional
- F05D2250/18—Two-dimensional patterned
- F05D2250/182—Two-dimensional patterned crenellated, notched
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2250/00—Geometry
- F05D2250/10—Two-dimensional
- F05D2250/18—Two-dimensional patterned
- F05D2250/184—Two-dimensional patterned sinusoidal
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2250/00—Geometry
- F05D2250/20—Three-dimensional
- F05D2250/29—Three-dimensional machined; miscellaneous
- F05D2250/294—Three-dimensional machined; miscellaneous grooved
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/20—Heat transfer, e.g. cooling
- F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2300/00—Materials; Properties thereof
- F05D2300/60—Properties or characteristics given to material by treatment or manufacturing
- F05D2300/603—Composites; e.g. fibre-reinforced
- F05D2300/6033—Ceramic matrix composites [CMC]
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Definitions
- This disclosure relates to sealing for adjacent components of a gas turbine engine.
- a gas turbine engine typically includes at least a compressor section, a combustor section and a turbine section.
- the compressor section pressurizes air into the combustion section where the air is mixed with fuel and ignited to generate an exhaust gas flow.
- the exhaust gas flow expands through the turbine section to drive the compressor section and, if the engine is designed for propulsion, a fan section.
- the turbine section may include multiple stages of rotatable blades and static vanes.
- An annular shroud or blade outer air seal may be provided around the blades in close radial proximity to the tips of the blades to reduce the amount of gas flow that escapes around the blades.
- the shroud typically includes a plurality of arc segments that are circumferentially arranged in an array. The arc segments are exposed to relatively hot gases in the gas flow path and may be configured to receive cooling airflow to cool portions of the shrouds.
- US 8 814 507 B1 discloses a cooling system for a three hook ring segment, comprising a blade outer air seal including an elongated seal body having a seal face
- US 2008/0131260 A1 discloses a method and system to facilitate cooling turbine engines
- WO 2015/138027 A2 discloses a meter plate for a blade outer air seal
- US 2017/175637 A1 discloses staged fuel and air injection in combustion systems of gas turbines.
- a seal assembly for a gas turbine engine according to claim 1.
- the first region defines a first volume.
- the second region defines a second volume, and the first volume is less than half of the second volume.
- the first section interconnects the second section and one or more inlet flow apertures defined by the seal body.
- the second section interconnects the first section and one or more outlet flow apertures defined by the seal body.
- the baffle is substantially free of any impingement holes such that the first region is fluidly isolated from the second region between the one or more inlet flow apertures and the one or more outlet flow apertures.
- the first region includes a third section that extends transversely from the first section such that the third section has a component in the axial direction (and may predominantly extend therein) to interconnect the first section and the one or more inlet flow apertures, with the third section extending along walls of the seal body that define the impingement face such that the first region at least partially surrounds the second region.
- a width of the first region varies at locations along the baffle in at least one of the circumferential direction and the axial direction.
- the first section extends along walls of a leading edge region of the seal body that defines the internal cavity.
- a pair of mounting blocks are insertable into respective openings along the mate faces to secure the blade outer air seal to an engine static structure.
- the first region includes outlet flow apertures along the mate faces.
- the seal is made of a first material including a ceramic material, and the seal has a unitary construction.
- the first region includes a third section that extends transversely from the first section such that the third section has a component in the axial direction (and may predominantly extend therein) to interconnect the first section and the at least one inlet flow aperture.
- the third section extends along walls of the seal body that define the impingement face such that the first region at least partially surrounds the second region.
- an end of the baffle abuts against the walls of the seal body defining the impingement face.
- the step of ejecting the cooling flow includes ejecting the cooling flow from outlet flow apertures of the first region along the mate faces.
- the engine static structure is an engine case that defines a plurality of cooling passages, and each of the plurality of cooling passages defines a passage axis that is oriented such that a projection of the passage axis intersects the impingement face.
- FIG. 1 schematically illustrates a gas turbine engine 20.
- the gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28.
- the fan section 22 drives air along a bypass flow path B in a bypass duct defined within a nacelle 15, while the compressor section 24 drives air along a core flow path C for compression and communication into the combustor section 26 then expansion through the turbine section 28.
- FIG. 1 schematically illustrates a gas turbine engine 20.
- the gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28.
- the fan section 22 drives air along a bypass flow path B in a bypass duct defined within a nacelle 15, while the compressor section 24 drives air along a core flow path C for compression and communication into the combustor section 26 then expansion through the turbine section 28.
- the exemplary engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided, and the location of bearing systems 38 may be varied as appropriate to the application.
- the low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a first (or low) pressure compressor 44 and a first (or low) pressure turbine 46.
- the inner shaft 40 is connected to the fan 42 through a speed change mechanism, which in exemplary gas turbine engine 20 is illustrated as a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30.
- the high speed spool 32 includes an outer shaft 50 that interconnects a second (or high) pressure compressor 52 and a second (or high) pressure turbine 54.
- a combustor 56 is arranged in exemplary gas turbine 20 between the high pressure compressor 52 and the high pressure turbine 54.
- a mid-turbine frame 57 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46.
- the mid-turbine frame 57 further supports bearing systems 38 in the turbine section 28.
- the inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.
- the core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded over the high pressure turbine 54 and low pressure turbine 46.
- the mid-turbine frame 57 includes airfoils 59 which are in the core airflow path C.
- the turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion.
- gear system 48 may be located aft of combustor section 26 or even aft of turbine section 28, and fan section 22 may be positioned forward or aft of the location of gear system 48.
- the engine 20 in one example is a high-bypass geared aircraft engine.
- the engine 20 bypass ratio is greater than about six, with an example embodiment being greater than about ten
- the geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine 46 has a pressure ratio that is greater than about five.
- the engine 20 bypass ratio is greater than about ten
- the fan diameter is significantly larger than that of the low pressure compressor 44
- the low pressure turbine 46 has a pressure ratio that is greater than about five.
- Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle.
- the geared architecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.
- the fan section 22 of the engine 20 is designed for a particular flight condition -- typically cruise at about 0.8 Mach and about 35,000 feet (10,668 m).
- the flight condition of 0.8 Mach and 35,000 ft (10,668 m), with the engine at its best fuel consumption - also known as "bucket cruise Thrust Specific Fuel Consumption ('TSFC')" - is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point.
- "Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system.
- the low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45.
- the "Low corrected fan tip speed" as disclosed herein according to one non-limiting embodiment is less than about 1150 ft / second (350.52 m/s).
- FIG. 2 shows selected portions of the turbine section 28 including a rotor 60 carrying one or more blades or airfoils 61 that are rotatable about the engine axis A.
- like reference numerals designate like elements where appropriate and reference numerals with the addition of one-hundred or multiples thereof designate modified elements that are understood to incorporate the same features and benefits of the corresponding original elements.
- Each airfoil 61 includes a platform 62 and an airfoil section 65 extending in a radial direction R from the platform 62 to a tip 64.
- the airfoil section 65 generally extends in a chordwise or axial direction X between a leading edge 66 and a trailing edge 68.
- a root section 67 of the airfoil 61 is mounted to, or integrally formed with, the rotor 60.
- a blade outer air seal (BOAS) 69 is spaced radially outward from the tip 64 of the airfoil section 65.
- the BOAS 69 can include a plurality of seal arc segments (one shown in Figure 5 at 169) that are circumferentially arranged in an annulus around the engine axis A.
- An array of the BOAS 69 are distributed about an array of the airfoils 61 to bound the core flow path C.
- a vane 70 is positioned along the engine axis A and adjacent to the airfoil 61.
- the vane 70 includes an airfoil section 71 extending between an inner platform 72 and an outer platform 73 to define a portion of the core flow path C.
- the turbine section 28 includes an array of airfoils 61, vanes 70, and BOAS 69 arranged circumferentially about the engine axis A.
- One or more cooling sources 75 are configured to provide cooling air to one or more cooling cavities or plenums 74 defined by an engine static structure such as the engine case 37 or another portion of the engine static structure 36 ( Figure 1 ).
- the engine case 37 extends along the engine axis A.
- the plenums 74 are defined between an engine case 37 and the outer platform 73 and/or BOAS 69.
- the plenums 74 are configured to receive pressurized cooling flow from the cooling source(s) 75 to cool portions of the airfoil 61, BOAS 69 and/or vane 70.
- Cooling sources 75 can include bleed air from an upstream stage of the compressor section 24 ( Figure 1 ), bypass air, or a secondary cooling system aboard the aircraft, for example.
- Each of the plenums 74 can extend in a circumferential or thickness direction T between adjacent airfoils 61, BOAS 69 and/or vanes 70.
- the tips 64 of each of the airfoil sections 65 and adjacent BOAS 69 are in close radial proximity to reduce the amount of gas flow that is redirected toward and over the rotating blade airfoil tips 64 through a corresponding clearance gap.
- Figures 3-9 illustrate an exemplary outer air seal case and support assembly 176 for sealing portions a gas turbine engine.
- the outer air seal case and support assembly 176 can be utilized for the seal assembly 76 of Figure 2 or incorporated into a portion of the engine 20 of Figure 1 , for example.
- the outer air seal 169 is a blade outer air seal (BOAS).
- Figure 3 is a sectional view of the outer air seal case and support assembly 176 in an installed position.
- Figure 4 is an axial view of adjacent outer air seal case and support assemblies 176 (indicated as 176A, 176B, 176C), and including a space eater baffle 190 installed in the outer air seal 169 (shown in dashed lines).
- Figure 5 illustrates a perspective view of the outer air seal 169.
- Figure 6 illustrates a support or mounting block 180 of the outer air seal case and support assembly 176 inserted into the outer air seal 169.
- Figure 7 illustrates a sectional view of the mounting block 180 between two adjacent outer air seals 169 (indicated as 169A, 169B).
- Figure 8 illustrates a sectional view of the outer air seal 169 with the space eater baffle 190 installed in the outer air seal 169.
- Figure 9 illustrates an isolated view of the example space eater baffle 190.
- baffles described herein may be utilized for other applications related to inner diameter (ID) and outer diameter (OD) endwall cooling, as well as airfoil convective and/or convective-film cooled cooling design configurations, for example.
- Utilization of the baffle arrangements disclosed herein may be incorporated for any cooling design concept that requires cooling air flow redistribution and tailored convective cooling as well as total and static pressure management to provide outlet flow apertures with the required pressure ratio for improvements in both internal convective heat transfer and film cooling performance to mitigate in-plane and thru wall temperature gradients.
- each outer air seal case & support assembly 176 includes an outer air seal 169, at least one support or mounting block 180 and at least one baffle 190 (shown in dashed lines in Figure 4 ).
- Each outer air seal 169 is arranged in close proximity to an airfoil tip 164 during operation of the engine.
- An array of the outer air seals 169 is circumferentially distributed about axis A and about an array of blades or airfoils 161 to bound a core flow path C (three seals 169A-169C shown in Figure 4 for illustrative purposes).
- Each outer air seal 169 includes an elongated main (or seal) body 170 that extends in a circumferential direction T between opposed (or first and second) mate faces 178 which define the bounds of the intersegment gaps between adjacent outer air seals 169 (see, e.g., intersegment gap G of Figure 7 ) and extends in an axial direction X between a leading edge portion 193 and a trailing edge portion 194.
- the main body 170 can have a generally elongated and arcuate profile, as illustrated by Figures 4 and 5 .
- the outer air seal 169 includes an inner diameter (ID) sealing surface portion 177 that extends circumferentially between the mate faces 178.
- the sealing surface portion 177 includes a front side or seal face surface 183 that extends circumferentially between the mate faces 178.
- the seal face surface 183 is oriented toward the engine axis A and bounds a gas path, such as the core flow path C, when the outer air seal 169 is located in an installed position.
- the sealing surface portion 177 includes a backside or impingement face 189 that is opposite to the seal face surface 183.
- the main body 170 extends in a radial direction R between the seal face surface 183 and impingement face 189.
- Each outer air seal 169 includes an engagement portion 179 that extends between the mate faces 178.
- the engagement portion 179 extends radially outward from the sealing surface portion 177 when in an installed position.
- the leading and trailing edge portions 193, 194 include the radially extending walls of the main body 170 along the sealing surface portion 177 and/or the engagement portion 179 that span between the mate faces 178.
- the outer air seal 169 includes an internal cavity 184 defined by the main body 170.
- the internal cavity 184 extends inwardly from at least one or a pair of openings 185 along each of the respective mate faces 178 (shown in Figure 5 ). In the illustrated example of Figures 3-8 , the internal cavity 184 extends circumferentially between the mate faces 178 and is defined between the sealing surface portion 177 and the engagement portion 179.
- the mounting block 180 can be arranged to secure one or more of the outer air seals 169 to a housing such as engine case 137, or to another portion of an engine static structure.
- An adjacent pair of outer air seals 169 are indicated as seals 169A, 169B in Figure 7 .
- the mounting block 180 includes at least one interface surface portion 181 extending outwardly from a main body or mounting portion 182.
- the mounting block 180 includes a pair of opposed interface surface portions 181A, 181B that extend in a direction predominately radially and circumferentially outward from the mounting portion 182.
- Each interface surface portion 181 is dimensioned to abut the engagement portion 179 of the respective seal 169 to limit relative movement in the radial and/or circumferential directions R, T, for example.
- a cross section of the mounting block 180 can have a generally trapezoidal geometry, as illustrated by Figures 6 and 7 .
- Surfaces of each interface surface portion 181 slope outwardly between a top 182A and bottom 182B of the mounting portion 182.
- the interface portions 181 can have a dovetail geometry.
- Each interface surface portion 181 can be inserted into or otherwise extend through a respective opening 185 to mate with ramped surfaces 186 of the internal cavity 184 (indicated as 186A, 186B in Figure 7 ) to bound movement of the seal 169 relative to the mounting block 180.
- the dovetail geometry circumferentially overlaps with walls of the engagement portions 179A, 179B when in the installed position to secure adjacent pairs of the outer air seals 169A, 169B to the engine case 137.
- Ends of the interface surface portions 181 can be contoured to guide the interface surface portions 181 through one of the openings 185 and into the respective internal cavity 184 during installation.
- the dovetail geometry and contouring can reduce mechanical stress on the outer air seal 169, including seals made of a composite material which can be strong but relative brittle.
- a pair of mounting blocks 180 are insertable into respective openings 185 along the mate faces 178 to secure the outer air seal 169 to the engine static structure 136, as illustrated by seal 169A of Figure 4 .
- Each interface surface portion 181 can include an outwardly extending retention feature 187.
- the retention feature 187 is dimensioned to abut against surfaces of the engagement portion 179 to seat the outer air seal 169 during assembly and limit circumferential and/or radial movement.
- the mounting block 180 can be secured to the engine case 137 using one or more fasteners 163 (one shown in Figure 7 for illustrative purposes).
- Each mounting portion 182 defines an aperture 188 that receives a respective fastener 163 to mechanically attach the mounting portion 182 to the engine case 137 and limit relative movement of one or more outer air seals 169.
- the fastener 163 is a bolt, and the aperture 188 threadably receives a length of the bolt.
- the fastener 163 is a clip or another structure to secure the outer air seal 169 to the engine static structure 136.
- the adjacent outer air seals 169A, 169B are arranged in close proximity such that the respective mate faces 178A, 178B define an intersegment gap G that extends a distance in the circumferential direction T.
- mate faces 578A, 578B of adjacent seals 569A, 569B are dimensioned to establish a ship lap joint to substantially reduce or eliminate the intersegment gap.
- the mounting block 180 is situated between the mate faces 178A, 178B such that the mounting block 180 spans across the intersegment gap G.
- a portion of the fastener 163 can be circumferentially aligned with one or more of the adjacent mate faces 178A, 178B and/or the intersegment gap G.
- the interface surface portions 181 abut against the adjacent outer air seals 169A, 169B to support the adjacent outer air seals 169A, 169B relative to the engine case 137 and limit circumferential and/or radial movement of the adjacent outer air seals 169A, 169B relative to the engine axis A.
- the mounting block 180 is arranged between the engagement portions 179A, 179B to circumferentially space apart the outer air seals 169A, 169B. Each mounting block 180 secures the engagement portions 179A, 179B to the engine case 137 when in the installed position. In alternative examples, the mounting block 180 is positioned at another location than along the intersegment gap G to secure the outer air seal 169 to the engine case 137, such as a midspan of the seal 169.
- the outer air seal assembly 176 defines a cooling arrangement 192 to deliver cooling flow F to portions of the outer air seal 169 and/or other portions of the outer air seal assembly 176 or components of the engine.
- the engine case 137 or another portion of the engine static structure 136 defines at least one or a plurality of cooling passage apertures 139.
- Plenum(s) 174 extend between the impingement face 189 and the engine case 137.
- the cooling passage apertures 139 are configured to communicate with and receive cooling flow from cooling source(s) 175.
- Each of the cooling passage apertures 139 defines a respective passage axis PA that is oriented such that a projection of the passage axis PA intersects the main body 170 of the outer air seal 169.
- the cooling passage apertures 139 eject cooling flow F into the plenum(s) 174 and toward the main body 170 of the outer air seal 169 in a direction along the passage axis PA.
- the main body 170 of the outer air seal 169 defines one or more inlet flow apertures 191 to the internal cavity 184.
- the inlet flow apertures 191 deliver or otherwise communicate cooling flow F from the cooling source 175 to the internal cavity 184 to cool adjacent portions of the seal 169 during engine operation.
- the impingement face 189 defines inlets of the inlet flow apertures 191 to serve as conduits to provide cooling flow F into the internal cavity 184.
- the inlets of the inlet flow apertures 191 may be distributed uniformly and/or non-uniformly in both and/or either the axial and/or circumferential directions X, T, for example.
- the inlet flow apertures 191 may vary in cross sectional area in either the axial and/or circumferential streamwise directions in order to optimize the distribution of cooling flow F, within the internal cavity 184, to achieve desired internal cavity Mach Number, Reynolds Number, and convective heat transfer, in order to tailor heat pickup and thermal cooling effectiveness requirements. Additionally the orientation and shape of the inlet flow apertures 191 may also comprise of both orthogonal and non-orthogonal apertures. In such instances the inlet flow apertures 191 may take the form of alternative geometries that may improve both thermal cooling and structural characteristics of the outer air seal 169. Inlet flow apertures 191 may consist of alternate geometric shapes and sizes, which may be, but are not limited to, conical, elliptical, and/or trapezoidal shapes of constant and/or varying cross sectional flow area and/or area ratio.
- One or more of the inlet flow apertures 191 can be spaced apart from each end wall of the engagement portion 179 that bounds the internal cavity 184.
- Each of the inlet flow apertures 191 defines an aperture axis AA ( Figure 3 ) such that a projection of the aperture axis AA intersects a wall of the internal cavity 184.
- the inlet flow apertures 191 are spaced from outlets of the cooling passage apertures 139.
- the inlet flow apertures 191 are substantially exposed and free of any obstructions when in an assembled position such that the inlet flow apertures 191 communicate cooling flow F to the internal cavity 184 during engine operation.
- Cooling flow F communicated by the inlet flow apertures 191 to the internal cavity 184 can reduce relatively sharp and/or large thermal gradients resulting in locally high thermal strains that may otherwise occur in the main body 170 due to a lack of active flow or cooling to the internal cavity 184. Reducing the potentially large radial and/or circumferential thermal gradients can mitigate detrimental thermal mechanical fatigue cracking resulting from high thermal-mechanical strains, which can improve the durability of outer air seals 169 comprising a ceramic or CMC material, for example.
- the inlet flow apertures 191 can be dimensioned to meter cooling flow F from the cooling source 175 into the internal cavity 184 at a desired rate and reduce inefficiencies due to excess cooling flow F that may otherwise be communicated from the cooling source 175 to provide localized convective cooling in addition to a more optimal distribution of pressures within internal cavity 184.
- the ability to effectively segregate, redirect, and redistribute internal flow and pressure within internal cavity 184 can improve the utilization of available cooling flow F to achieve leakage and/or film cooling outflow requirements that may be needed to prevent ingestion and/or entrainment of the hot gases of the core flow path C.
- the tailoring and redistribution of cooling flow F can provide for a more efficient outer air seal 169 thermal cooling design configuration. By effectively utilizing less cooling flow F, improvements in thermodynamic cycle performance and outer air seal 169 durability capability can be achieved.
- Inlet flow apertures 191 may be used to locally tailor the distribution of cooling flow F to both axial and circumferntial portions of internal cavity 184. Additionally, the local internal pressure within internal cavity 184 may be controlled and tailored to provide desired pressure ratio requirements to outlet flow apertures 197, 197" (shown in Figure 8 ) and/or exposed mate faces 178 at either end of the outer air seal 169 (shown in Figure 6 ) in order to prevent ingestion and/or entrainment of the hot gases in the core flow path C. It may be desirable to keep the cooling flow F active and pressurized within cavity 184 in order to prevent contaminants and other environmental particulate from depositing itself in an otherwise stagnate cavity which may adversely impact the thermal mechanical characteristics of the material properties of the outer air seal 169. Unique tailoring of the inlet flow apertures 191 also can provide additional mitigation of local hot spots and sharp thermal gradients induced by variations in both local circumferential and axial heat flux and temperature distributions along the ID sealing surface portion 177.
- the inlet flow apertures 191 can have a generally elliptical cross sectional profile and extend in a radial and/or predominately radial direction R through a thickness of a wall of the main body 170. Additionally, inlet flow apertures 191 may be non-orthogonal and oriented in both a circumferential and/or axial direction. Alternative inlet flow apertures 191 geometries may be utilized either independently and/or in conjunction with, including, but not limited to, slots, conical, diffused, single or multi-lobed, shapes and/or consist of more complex geometric shapes comprising of single or multiple concave and/or convex surfaces.
- the inlet flow apertures 191 can be arranged relative to the cooling passage apertures 139 to provide localized cooling to selected portions of the outer air seal 169.
- the projection of the passage axis PA of at least some cooling passage apertures 139 can intersect the impingement face 189 to provide localized impingement cooling, but the projection of the passage axis PA of other cooling passage apertures 139 may not.
- Each passage axis PA can be perpendicular or transverse to the impingement face 189 at the point of intersection (see, e.g., passage apertures 139A and 139B of Figure 3 , respectively).
- the passage axis PA of at least some of the cooling passage apertures 139 can be aligned with or otherwise intersect a respective one of the inlets 191 (see, e.g., passage apertures 139A) to increase an impingement distance between an outlet of the respective passage apertures 139 and the seal 169.
- the increased impingement distance can reduce the heat transfer augmentation, which can reduce thermal gradients and more uniformly distribute the cooling flow F in the radial direction R.
- Some of the inlet flow apertures 191 can be offset from a projection of the passage axis PA of the cooling passage apertures 139, as illustrated by inlets 191' in Figure 3 (shown in dashed lines). Offsetting the inlet flow apertures 191' allows the cooling flow F to follow along surfaces of the impingement face 189 to cool adjacent portions of the outer air seal 169 prior to entry into the inlet flow apertures 191'. Aligning some of the cooling passage apertures 139 with respective inlet flow apertures 191 but offsetting other cooling passage apertures 139 with respect to each of the inlet flow apertures 191 can provide a mixture of localized, impingement cooling to surfaces of the impingement face 189 and more direct cooling to the internal cavity 184. In examples, each inlet flow aperture 191 is aligned with a respective cooling passage apertures 139. In other examples, each inlet flow aperture 191 is offset from each cooling passage apertures 139.
- the outer air seal 169 is exposed to temperature differentials between the gases in the core flow path C and the cooling flow F from the cooling source 175. Portions of the outer air seal 169 may be exposed to different thermal loads, which can result in thermal gradients across the outer air seal 169. The thermal gradients may cause cracks to form in the outer air seal 169, for example.
- the baffle 190 is secured to the seal 169 in at least a portion of the internal cavity 184.
- Figure 3 illustrates the seal assembly 176 with the baffle 190 uninstalled in the internal cavity 184.
- the baffle 190 can be dimensioned to be removably inserted into the internal cavity 184.
- the baffle 190 is permanently secured to in the internal cavity 184, such as by molding the baffle 190 in the main body 170, for example.
- the baffle 190 divides the internal cavity 184 into a plurality of localized regions.
- the baffle 190 divides the internal cavity 184 into at least a first region 195 and a second region 196 that is separate and distinct from the first region 195.
- the first region 195 serves as a cooling channel to direct or otherwise guide the cooling flow F along a perimeter of the internal cavity 184.
- the first region 195 can provide an increase in the max flux as well as optimized convective heat transfer coefficient (HTC) augmentation to radially inner portions of the internal cavity 184 adjacent to the seal face surface 183.
- HTC convective heat transfer coefficient
- the first region 195 can be established along portions of the outer air seal 169 that may be more susceptible to thermal loads or gradients, thereby establishing a prioritized cooling scheme.
- the prioritized cooling can reduce thermal gradients in the outer air seal 169, thereby reducing a likelihood of cracking and improving durability.
- the regions 195, 196 can be dimensioned to provide localized cooling to adjacent portions of the outer air seal 169 with respect to expected heat loads and thermal gradients that may be experienced during engine operation.
- the first region 195 defines a first volume
- the second region 196 defines a second volume
- the first volume is less than half of the second volume.
- the first volume is less than 10% or 25% of the second volume.
- At least one inlet flow apertures 191 is defined along the first region 195.
- the inlet flow apertures 191 can be defined along the impingement face 189, for example.
- the first region 195 communicates cooling flow F from the inlet flow apertures 191 to one or more outlet flow apertures 197 defined by the main body 170.
- one or more outlet flow apertures 197 are defined along the trailing edge portion 194 of the main body 170.
- the sealing surface portion 177 defines one or more outlet flow apertures 197' (shown in dashed lines) that serve as film cooling holes to eject cooling flow F from the internal cavity 184 into the core flow path C to cool adjacent surfaces of the outer air seal face surface 183.
- one or more outlet flow apertures 197" are defined along the leading edge portion 193 of the seal 169. The outlet flow apertures 197" can reduce thermal gradients across a thickness of the walls of the leading edge portion 193.
- the baffle 190 is arranged in the internal cavity 184 to change a direction of the cooling flow F from the inlet flow aperture 191 relative to the respective aperture axis AA.
- the baffle 190 includes an elongated baffle body 190-1 having a generally C-shaped geometry.
- a first end 190-2 of the baffle 190 is contoured and is dimensioned to abut against surfaces of the internal cavity 184 along walls of the main body 170 that define the impingement face 189.
- the baffle 190 fluidly isolates the first and second regions 195, 196 along the first end 190-2.
- a second end 190-3 of the baffle 190 can be substantially planar.
- second end 190-3' (shown in dashed lines in Figure 9 ) is contoured and is dimensioned to abut against surfaces of the internal cavity 184 such that the baffle 190 fluidly isolates the first and second regions 195, 196 along the second end 190-3'.
- the baffle 190 is substantially free of any impingement holes such that the first region 195 is fluidly isolated from the second region 196 between the inlet flow apertures 191 and the outlet flow apertures 197.
- the term "substantially free” means that less than 3% of the baffle volume defines holes.
- the baffle 190 includes one or more impingement holes 199 (shown in dashed lines in Figure 8 ).
- outlet flow apertures 199 described herein are referred to as impingement holes it is to be understood that these features may consist of noncircular and cylindrical shapes, such as but not limited to, slot, elliptical, oval, race track, or other multi-faceted geometry shapes comprising of conical and/or convex surfaces.
- ends of the baffle 190 are spaced apart from walls of the internal cavity 184 such that the first region 195 and/or the second region 196 includes outlets 198 along the mate faces 178.
- the baffle 190 can include one or more bumpers or protrusions 190-4 ( Figure 9 ) such as pedestals to space apart the baffle 190 and offset it from the surfaces of the internal cavity 184, creating a gap or channel by which cooling flow F may be regulated and/or controlled and then dispersed through outlet flow apertures 197, 197' and 197" illustrated in Figure 8 .
- the mounting block 180 can be dimensioned relative to surfaces of the internal cavity 184 such that cooling flow F in the internal cavity 184 can exit from the outlets 198, into openings 185 along the mate faces 178, and into the intersegment gaps G. Communication of the cooling flow F from the outlets 198 can cool portions of the mate faces 178, thereby improving durability of the outer air seals 169. Communicating the cooling flow F into each intersegment gap G creates a fluidic sealing boundary and cooling purge flow relationship to be formed, which can reduce the likelihood of ingestion of hot combustion gases from the core flow path C and into the intersegment gap G.
- ends of the baffle 190 are dimensioned to abut against the walls of the internal cavity 184 such that the first region 195 is substantially fluidly isolated from the mate faces 178.
- the baffle 190 can be dimensioned such that the first region 195 has a circuitous path between surfaces of the baffle 190 and walls of the main body 170 that define the internal cavity 184.
- the baffle 190 directs or guides the cooling flow F from the inlet flow apertures 191 along the circuitous path to provide localized cooling to adjacent portions of the main body 170.
- the first region 195 includes a circuitous path that follows a perimeter of the internal cavity 184.
- the circuitous path of the first region is defined by at least first, second and third sections 195-1, 195-2, 195-3.
- the baffle 190 can be arranged to establish a relatively small offset from walls of the internal cavity 184 to improve heat transfer augmentation to portions of the seal 169 adjacent to the first, second and third sections 195-1, 195-2, 195-3.
- an offset or width W1 of the baffle 190 from surfaces of the internal cavity 184 along the first region 195 can be less than 5% or 10% of a height H and/or length L of the internal cavity 184.
- the relatively small offset can increase velocities of the cooling flow F along the first region 195.
- the width W1 along the first region 195 can vary for each of the first, second and third sections 195-1, 195-2, 195-3.
- the varying offset along the first region 195 can change the Mach No. distribution of the cooling flow along the flow path to provide prioritized cooling augmentation to surfaces of the seal 169 along the first region 195.
- width W1-1 along the first section 195-1 can be greater than width W1-2 along the second section 195-2 to cause relatively greater cooling flow F to circulate along the first section 195-1.
- the relatively greater width W1-1 can reduce a velocity of the cooling flow F, thereby increasing the local static pressure within the channel between the baffle 190 and the wall surfaces of internal cavity 184, along the first section 195-1, and the relatively lesser width W1-2 can the decrease a static pressure of the cooling flow F to reduce the pressure ratio across the outlet flow apertures 197, 197' for improved film cooling.
- the gap between the baffle 190 and the wall surfaces of internal cavity 184 may be set by altering the channel heights W1-1 and W1-2 in order to achieve the desired local cooling flow F velocity and required local static pressure to ensure sufficient driving pressure exists across outlet flow apertures 197, 197', and 197" in order to achieve the desired film cooling and backside convective heat transfer characteristics in order to reduce local temperatures, minimize thermal gradients, and increase the local film cooling performance of the out air seal 169.
- the first section 195-1 extends transversely from the second and third sections 195-2, 195-3 such that a cross section of the first region 195 has a generally C-shaped geometry.
- the first section 195-1 has a component in the radial direction R with respect to the engine axis A, for example.
- the first section 195-1 extends along walls of the leading edge portion 193 that define the internal cavity 184.
- the baffle 190 can be situated in other orientations.
- an orientation of the baffle 190 can be inverse to the orientation of the baffle 190 illustrated in Figure 8 such that the first section 195-1 extends along walls of the trailing edge portion 194 that define the internal cavity 184.
- the second section 195-2 has a component in the axial direction X with respect to the engine axis A, for example, such that the second section 195-2 is defined between the baffle 190 and walls of the main body 170 defining the seal face surface 183.
- the third section 195-3 has a component in the axial direction X with respect to the engine axis A, for example, such that the third section 195-3 is defined between the baffle 190 and walls of the main body 170 defining the impingement face 189.
- the third section 195-3 fluidically connects the first section 195-1 and the inlets 191.
- the third section 195-3 extends along walls of the main body 170 that define the impingement face 189 such that the first region 195 at least partially surrounds the second region 196.
- the second region 196 extends along sidewalls of the trailing edge portion 194, but is spaced apart from walls of the leading edge portion 193.
- a length of the third section 195-3 can be less than, equal to, or greater than a length of the second section 195-2 with respect to the axial direction X or engine axis A depending on flow, heat transfer and pressure loss requirements needed to minimize local thermal gradients while ensure positive pressure ratio and outflow conditions for outlet flow apertures 197, 197', and 197" as shown in Figure 8 .
- the location and number of flow apertures can be selected depending on outer air seal 169 cooling flow F allocations and durability life requirements.
- the first section 195-1 fluidically connects the second section 195-2 and the inlet flow apertures 191 to deliver cooling flow F from the cooling source 175 to the second section 195-2.
- the second section 195-2 fluidically connects the first section 195-1 and one or more outlets 197 defined by the main body 170 to deliver the cooling flow F to downstream portions of the outer air seal 169.
- the cooling flow F is ejected from outlets 198 of the first region 195 along the mate faces 178, as illustrated by Figure 7 .
- the second end 190-3 of the baffle 190 is spaced apart from walls of the internal cavity 184 such that cooling flow F is communicated from the second section 195-2 to the second region 196 subsequent to be guided along the circuitous path established by the first region 195.
- the outer air seal 169 is made of a first material
- the mounting block 180 and/or baffle 190 is made of a second, different material.
- the first material can include a ceramic or ceramic matrix composite (CMC) material.
- CMC ceramic or ceramic matrix composite
- the outer air seal 169 can be formed from one or more layers L of a CMC layup ( Figure 3 ).
- the outer air seal 169 can be made of another material, such as a high temperature metal, alloy, or composite material.
- the mounting block 180 and/or baffle 190 can be made of a second material such as a high temperature composite, metal, or alloy, such as a nickel or cobalt-based superalloy, for example.
- the baffle 190 can include one or more coatings, such as to reduce interaction between the CMC material of the outer air seal 169 and the nickel alloy material of the baffle 190.
- the first and second materials can differ.
- the outer air seal 169 is made of a first material
- the mounting block 180 and/or baffle 190 is made of a second material that is the same as the first material, including any of the materials disclosed herein.
- the outer air seal 169 is formed to have a unitary construction.
- the sealing surface portion 177 and each engagement portion 179 are separate and distinct components that are mechanically attached to one another with one or more fasteners.
- the baffle 190 is made from sheet metal, or other materials with moderate temperature capability, which may include but not limited to Inconel, Cobalt-Chrome, Hastelloy, Nickel etc. and is formed to the desired geometry shape.
- the inlet and outlet flow apertures 191, 197 can be formed by a machining operation after forming the main body 170. Alternatively, the inlet and outlet apertures 191, 197 can be formed during the fabrication of the main body 170 by arranging the layers L of the CMC layup ( Figure 3 ), for example.
- a method of sealing is as follows. With reference to Figures 7 and 8 , the baffle 190 is inserted into an opening 185 to the internal cavity 184 such that the baffle 190 is secured to the outer air seal 169.
- a mounting block 180 is positioned between mate faces 178A, 178B of each adjacent pair of seals 169A, 169B.
- Each of the interface surface portions 181A, 181B is inserted or otherwise moved in a circumferential direction T through a respective one of the openings 185A, 185B and into abutment with the ramp surfaces 186A, 186B of the engagement portions 179A, 179B to bound movement of the outer air seals 169A, 169B.
- the fastener 163 is mechanically attached or otherwise secured to the mounting block 180 to secure the outer air seal assembly 176 to the engine static structure 136.
- pressurized cooling flow F is communicated from the cooling source 175 to the cooling passage apertures 139.
- the cooling passages apertures 139 eject the cooling flow F into the plenum 174 and in a direction toward the impingement face 189 of the seal 169.
- the cooling flow F can be discharged from the cooling passage apertures 139 at a predetermined pressure and velocity such that the cooling flow F impinges on localized surfaces of the impingement face 189.
- the cooling flow F circulates from the plenum 174 and into the inlet apertures 191.
- the cooling flow F circulates from the inlet flow apertures 191 and disperses into the third section 195-3 of the first region 195.
- the cooling flow F can circulate from the inlet aperture(s) 191 to the third section 195-3, then to the first section 195-1, and then to the second section 195-2.
- the cooling flow F can be ejected from one or more outlets 197, 197' along the first and/or second sections 195-1, 195-2 and into the gas path such as the core flow path C.
- the cooling flow F circulates from the plenum 174 into the inlet flow aperture(s) 191', and then into the second region 196.
- FIG 10 illustrates a baffle 290 according to another example.
- the baffle 290 can have a non-uniform offset from surfaces of the internal cavity 284.
- walls of baffle body 290-1 are contoured such that a distance or width of first region 295 between the baffle 290 and walls of main body 270 varies at locations along the first region 295 in a first direction D1, which can be at least one of the circumferential, radial and/or axial directions T, R, X, and/or a combination thereof (see Figures 3-4 and 8 ), and wherein a passage axis of the inlet cooling aperture(s) (see, e.g., passage axis PA of Figure 8 ) can be oriented orthogonally or non-orthogonally with respect to grooves 290-4 to minimize or otherwise reduce pressure losses, for example.
- the baffle body 290-1 defines one or more grooves 290-4 such that an offset or width W1 between surfaces of the grooves 290-4 and the main body 270 is greater than a width W2 between the main body 270 and other portions of the baffle body 290-1 along the first region 295.
- a cross section of the grooves 290-4 can have a substantially rectangular profile, as illustrated by Figure 10 .
- the grooves can have other geometries, such as a generally concave geometry as illustrated by grooves 390-4 of Figure 11 .
- the features of Figures 10 and 11 can be incorporated into the baffle 190 along the first, second and/or third sections 195-1, 195-2, 195-3 of the first region 195 of Figure 8 , for example.
- the grooves 290-4/390-4 can be circumferentially aligned with respect to injection nozzles of the combustor 56 or vanes 70 axially forward of the outer air seal 169, which may cause non-uniform thermal loads on the outer air seal 169, for example.
- FIG. 12 illustrates adjacent seal assemblies 476 according to another example.
- Two adjacent seals are labeled as 469A, 469B.
- Each baffle 490 is secured between a respective seal 469A/469B and surfaces of mounting block 480.
- baffle 490 is dimensioned such that portions of a baffle body 490-1 of the baffle 490, such as a portion of first end 490-2, are trapped between interface portions 481 of the mounting block 480 and ramped surfaces 486 of internal cavity 484 defined by the respective seal 469 that are opposed to surfaces of the respective interface portion 481 to limit relative movement in the axial, radial and/or circumferential directions X, R, T, for example.
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
- Gasket Seals (AREA)
Description
- This disclosure relates to sealing for adjacent components of a gas turbine engine.
- A gas turbine engine typically includes at least a compressor section, a combustor section and a turbine section. The compressor section pressurizes air into the combustion section where the air is mixed with fuel and ignited to generate an exhaust gas flow. The exhaust gas flow expands through the turbine section to drive the compressor section and, if the engine is designed for propulsion, a fan section.
- The turbine section may include multiple stages of rotatable blades and static vanes. An annular shroud or blade outer air seal may be provided around the blades in close radial proximity to the tips of the blades to reduce the amount of gas flow that escapes around the blades. The shroud typically includes a plurality of arc segments that are circumferentially arranged in an array. The arc segments are exposed to relatively hot gases in the gas flow path and may be configured to receive cooling airflow to cool portions of the shrouds.
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US 8 814 507 B1 discloses a cooling system for a three hook ring segment, comprising a blade outer air seal including an elongated seal body having a seal face,US 2008/0131260 A1 discloses a method and system to facilitate cooling turbine engines,WO 2015/138027 A2 discloses a meter plate for a blade outer air seal, andUS 2017/175637 A1 discloses staged fuel and air injection in combustion systems of gas turbines. - According to a first aspect of the invention there is provided a seal assembly for a gas turbine engine according to claim 1.
- In an embodiment, the first region defines a first volume. The second region defines a second volume, and the first volume is less than half of the second volume.
- In a further embodiment of any of the foregoing embodiments, the first section interconnects the second section and one or more inlet flow apertures defined by the seal body.
- In a further embodiment of any of the foregoing embodiments, the second section interconnects the first section and one or more outlet flow apertures defined by the seal body.
- In a further embodiment of any of the foregoing embodiments, the baffle is substantially free of any impingement holes such that the first region is fluidly isolated from the second region between the one or more inlet flow apertures and the one or more outlet flow apertures.
- In a further embodiment of any of the foregoing embodiments, the first region includes a third section that extends transversely from the first section such that the third section has a component in the axial direction (and may predominantly extend therein) to interconnect the first section and the one or more inlet flow apertures, with the third section extending along walls of the seal body that define the impingement face such that the first region at least partially surrounds the second region.
- In a further embodiment of any of the foregoing embodiments, a width of the first region varies at locations along the baffle in at least one of the circumferential direction and the axial direction.
- In a further embodiment of any of the foregoing embodiments, the first section extends along walls of a leading edge region of the seal body that defines the internal cavity.
- In a further embodiment of any of the foregoing embodiments, a pair of mounting blocks are insertable into respective openings along the mate faces to secure the blade outer air seal to an engine static structure.
- In a further embodiment of any of the foregoing embodiments, the first region includes outlet flow apertures along the mate faces.
- In a further embodiment of any of the foregoing embodiments, the seal is made of a first material including a ceramic material, and the seal has a unitary construction.
- According to another aspect of the invention there is provided a method of sealing of a gas turbine engine according to claim 13.
- In an embodiment, the first region includes a third section that extends transversely from the first section such that the third section has a component in the axial direction (and may predominantly extend therein) to interconnect the first section and the at least one inlet flow aperture. The third section extends along walls of the seal body that define the impingement face such that the first region at least partially surrounds the second region.
- In a further embodiment of any of the foregoing embodiments, an end of the baffle abuts against the walls of the seal body defining the impingement face.
- In a further embodiment of any of the foregoing embodiments, the step of ejecting the cooling flow includes ejecting the cooling flow from outlet flow apertures of the first region along the mate faces.
- In a further embodiment of any of the foregoing embodiments, the engine static structure is an engine case that defines a plurality of cooling passages, and each of the plurality of cooling passages defines a passage axis that is oriented such that a projection of the passage axis intersects the impingement face.
- The various features and advantages of this invention will become apparent to those skilled in the art from the following detailed description of an embodiment. The drawings that accompany the detailed description can be briefly described as follows.
- The various features and advantages of this invention will become apparent to those skilled in the art from the following detailed description of an embodiment. The drawings that accompany the detailed description can be briefly described as follows.
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Figure 1 shows a gas turbine engine. -
Figure 2 shows an airfoil arrangement for a turbine section. -
Figure 3 illustrates a sectional view of a seal assembly including a seal. -
Figure 4 illustrates an axial view of adjacent seal assemblies, including a baffle inserted into the seal ofFigure 3 . -
Figure 5 illustrates an isolated perspective view of the seal ofFigure 3 . -
Figure 6 illustrates a perspective view of a support inserted into the seal ofFigure 3 . -
Figure 7 illustrates a sectional view of the support ofFigure 6 between two adjacent seals. -
Figure 8 illustrates a sectional view of the seal along line 8-8 ofFigure 4 . -
Figure 9 illustrates an isolated perspective view of the baffle ofFigure 8 . -
Figure 10 illustrates a sectional view of a seal assembly according to another example. -
Figure 11 illustrates a sectional view of a seal assembly according to yet another example. -
Figure 12 illustrates a sectional view of a pair of baffles secured between a support adjacent seals. -
Figure 13 illustrates adjacent seal assemblies establishing a ship lap joint according to an example. -
Figure 1 schematically illustrates agas turbine engine 20. Thegas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates afan section 22, acompressor section 24, a combustor section 26 and aturbine section 28. Thefan section 22 drives air along a bypass flow path B in a bypass duct defined within anacelle 15, while thecompressor section 24 drives air along a core flow path C for compression and communication into the combustor section 26 then expansion through theturbine section 28. Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures. - The
exemplary engine 20 generally includes alow speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an enginestatic structure 36 viaseveral bearing systems 38. It should be understood thatvarious bearing systems 38 at various locations may alternatively or additionally be provided, and the location ofbearing systems 38 may be varied as appropriate to the application. - The
low speed spool 30 generally includes aninner shaft 40 that interconnects afan 42, a first (or low) pressure compressor 44 and a first (or low)pressure turbine 46. Theinner shaft 40 is connected to thefan 42 through a speed change mechanism, which in exemplarygas turbine engine 20 is illustrated as a gearedarchitecture 48 to drive thefan 42 at a lower speed than thelow speed spool 30. Thehigh speed spool 32 includes anouter shaft 50 that interconnects a second (or high)pressure compressor 52 and a second (or high)pressure turbine 54. Acombustor 56 is arranged inexemplary gas turbine 20 between thehigh pressure compressor 52 and thehigh pressure turbine 54. Amid-turbine frame 57 of the enginestatic structure 36 is arranged generally between thehigh pressure turbine 54 and thelow pressure turbine 46. Themid-turbine frame 57 further supports bearingsystems 38 in theturbine section 28. Theinner shaft 40 and theouter shaft 50 are concentric and rotate viabearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes. - The core airflow is compressed by the low pressure compressor 44 then the
high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over thehigh pressure turbine 54 andlow pressure turbine 46. Themid-turbine frame 57 includesairfoils 59 which are in the core airflow path C. Theturbines low speed spool 30 andhigh speed spool 32 in response to the expansion. It will be appreciated that each of the positions of thefan section 22,compressor section 24, combustor section 26,turbine section 28, and fandrive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, andfan section 22 may be positioned forward or aft of the location ofgear system 48. - The
engine 20 in one example is a high-bypass geared aircraft engine. In a further example, theengine 20 bypass ratio is greater than about six, with an example embodiment being greater than about ten, the gearedarchitecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and thelow pressure turbine 46 has a pressure ratio that is greater than about five. In one disclosed embodiment, theengine 20 bypass ratio is greater than about ten, the fan diameter is significantly larger than that of the low pressure compressor 44, and thelow pressure turbine 46 has a pressure ratio that is greater than about five.Low pressure turbine 46 pressure ratio is pressure measured prior to inlet oflow pressure turbine 46 as related to the pressure at the outlet of thelow pressure turbine 46 prior to an exhaust nozzle. The gearedarchitecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans. - A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The
fan section 22 of theengine 20 is designed for a particular flight condition -- typically cruise at about 0.8 Mach and about 35,000 feet (10,668 m). The flight condition of 0.8 Mach and 35,000 ft (10,668 m), with the engine at its best fuel consumption - also known as "bucket cruise Thrust Specific Fuel Consumption ('TSFC')" - is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. "Low fan pressure ratio" is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane ("FEGV") system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. "Low corrected fan tip speed" is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram °R) / (518.7 °R)]^0.5 (where °R = K x 9/5). The "Low corrected fan tip speed" as disclosed herein according to one non-limiting embodiment is less than about 1150 ft / second (350.52 m/s). -
Figure 2 shows selected portions of theturbine section 28 including arotor 60 carrying one or more blades orairfoils 61 that are rotatable about the engine axis A. In this disclosure, like reference numerals designate like elements where appropriate and reference numerals with the addition of one-hundred or multiples thereof designate modified elements that are understood to incorporate the same features and benefits of the corresponding original elements. Eachairfoil 61 includes aplatform 62 and anairfoil section 65 extending in a radial direction R from theplatform 62 to atip 64. Theairfoil section 65 generally extends in a chordwise or axial direction X between aleading edge 66 and a trailingedge 68. Aroot section 67 of theairfoil 61 is mounted to, or integrally formed with, therotor 60. A blade outer air seal (BOAS) 69 is spaced radially outward from thetip 64 of theairfoil section 65. TheBOAS 69 can include a plurality of seal arc segments (one shown inFigure 5 at 169) that are circumferentially arranged in an annulus around the engine axis A. An array of theBOAS 69 are distributed about an array of theairfoils 61 to bound the core flow path C. - A
vane 70 is positioned along the engine axis A and adjacent to theairfoil 61. Thevane 70 includes anairfoil section 71 extending between aninner platform 72 and anouter platform 73 to define a portion of the core flow path C. Theturbine section 28 includes an array ofairfoils 61,vanes 70, andBOAS 69 arranged circumferentially about the engine axis A. - One or more cooling sources 75 (one shown) are configured to provide cooling air to one or more cooling cavities or
plenums 74 defined by an engine static structure such as the engine case 37 or another portion of the engine static structure 36 (Figure 1 ). The engine case 37 extends along the engine axis A. In the illustrated example ofFigure 2 , theplenums 74 are defined between an engine case 37 and theouter platform 73 and/orBOAS 69. Theplenums 74 are configured to receive pressurized cooling flow from the cooling source(s) 75 to cool portions of theairfoil 61,BOAS 69 and/orvane 70.Cooling sources 75 can include bleed air from an upstream stage of the compressor section 24 (Figure 1 ), bypass air, or a secondary cooling system aboard the aircraft, for example. Each of theplenums 74 can extend in a circumferential or thickness direction T betweenadjacent airfoils 61,BOAS 69 and/orvanes 70. Thetips 64 of each of theairfoil sections 65 andadjacent BOAS 69 are in close radial proximity to reduce the amount of gas flow that is redirected toward and over the rotatingblade airfoil tips 64 through a corresponding clearance gap. -
Figures 3-9 illustrate an exemplary outer air seal case andsupport assembly 176 for sealing portions a gas turbine engine. The outer air seal case andsupport assembly 176 can be utilized for theseal assembly 76 ofFigure 2 or incorporated into a portion of theengine 20 ofFigure 1 , for example. In the illustrated example ofFigures 3-9 , theouter air seal 169 is a blade outer air seal (BOAS).Figure 3 is a sectional view of the outer air seal case andsupport assembly 176 in an installed position.Figure 4 is an axial view of adjacent outer air seal case and support assemblies 176 (indicated as 176A, 176B, 176C), and including aspace eater baffle 190 installed in the outer air seal 169 (shown in dashed lines).Figure 5 illustrates a perspective view of theouter air seal 169.Figure 6 illustrates a support or mountingblock 180 of the outer air seal case andsupport assembly 176 inserted into theouter air seal 169.Figure 7 illustrates a sectional view of the mountingblock 180 between two adjacent outer air seals 169 (indicated as 169A, 169B).Figure 8 illustrates a sectional view of theouter air seal 169 with thespace eater baffle 190 installed in theouter air seal 169.Figure 9 illustrates an isolated view of the examplespace eater baffle 190. The baffles described herein may be utilized for other applications related to inner diameter (ID) and outer diameter (OD) endwall cooling, as well as airfoil convective and/or convective-film cooled cooling design configurations, for example. Utilization of the baffle arrangements disclosed herein may be incorporated for any cooling design concept that requires cooling air flow redistribution and tailored convective cooling as well as total and static pressure management to provide outlet flow apertures with the required pressure ratio for improvements in both internal convective heat transfer and film cooling performance to mitigate in-plane and thru wall temperature gradients. - Referring to
Figures 3-5 , each outer air seal case &support assembly 176 includes anouter air seal 169, at least one support or mountingblock 180 and at least one baffle 190 (shown in dashed lines inFigure 4 ). Eachouter air seal 169 is arranged in close proximity to anairfoil tip 164 during operation of the engine. An array of the outer air seals 169 is circumferentially distributed about axis A and about an array of blades orairfoils 161 to bound a core flow path C (threeseals 169A-169C shown inFigure 4 for illustrative purposes). - Each
outer air seal 169 includes an elongated main (or seal)body 170 that extends in a circumferential direction T between opposed (or first and second) mate faces 178 which define the bounds of the intersegment gaps between adjacent outer air seals 169 (see, e.g., intersegment gap G ofFigure 7 ) and extends in an axial direction X between aleading edge portion 193 and a trailingedge portion 194. Themain body 170 can have a generally elongated and arcuate profile, as illustrated byFigures 4 and5 . Theouter air seal 169 includes an inner diameter (ID) sealingsurface portion 177 that extends circumferentially between the mate faces 178. The sealingsurface portion 177 includes a front side or sealface surface 183 that extends circumferentially between the mate faces 178. Theseal face surface 183 is oriented toward the engine axis A and bounds a gas path, such as the core flow path C, when theouter air seal 169 is located in an installed position. The sealingsurface portion 177 includes a backside orimpingement face 189 that is opposite to theseal face surface 183. Themain body 170 extends in a radial direction R between theseal face surface 183 andimpingement face 189. - Each
outer air seal 169 includes anengagement portion 179 that extends between the mate faces 178. Theengagement portion 179 extends radially outward from the sealingsurface portion 177 when in an installed position. The leading and trailingedge portions main body 170 along the sealingsurface portion 177 and/or theengagement portion 179 that span between the mate faces 178. - The
outer air seal 169 includes aninternal cavity 184 defined by themain body 170. Theinternal cavity 184 extends inwardly from at least one or a pair ofopenings 185 along each of the respective mate faces 178 (shown inFigure 5 ). In the illustrated example ofFigures 3-8 , theinternal cavity 184 extends circumferentially between the mate faces 178 and is defined between the sealingsurface portion 177 and theengagement portion 179. - Referring to
Figures 6 and 7 , the mountingblock 180 can be arranged to secure one or more of the outer air seals 169 to a housing such as engine case 137, or to another portion of an engine static structure. An adjacent pair of outer air seals 169 are indicated asseals Figure 7 . The mountingblock 180 includes at least oneinterface surface portion 181 extending outwardly from a main body or mountingportion 182. In the illustrated example ofFigures 6 and 7 , the mountingblock 180 includes a pair of opposedinterface surface portions portion 182. Eachinterface surface portion 181 is dimensioned to abut theengagement portion 179 of therespective seal 169 to limit relative movement in the radial and/or circumferential directions R, T, for example. - A cross section of the mounting
block 180 can have a generally trapezoidal geometry, as illustrated byFigures 6 and 7 . Surfaces of eachinterface surface portion 181 slope outwardly between a top 182A and bottom 182B of the mountingportion 182. Theinterface portions 181 can have a dovetail geometry. Eachinterface surface portion 181 can be inserted into or otherwise extend through arespective opening 185 to mate with rampedsurfaces 186 of the internal cavity 184 (indicated as 186A, 186B inFigure 7 ) to bound movement of theseal 169 relative to themounting block 180. The dovetail geometry circumferentially overlaps with walls of theengagement portions interface surface portions 181 can be contoured to guide theinterface surface portions 181 through one of theopenings 185 and into the respectiveinternal cavity 184 during installation. The dovetail geometry and contouring can reduce mechanical stress on theouter air seal 169, including seals made of a composite material which can be strong but relative brittle. A pair of mountingblocks 180 are insertable intorespective openings 185 along the mate faces 178 to secure theouter air seal 169 to the engine static structure 136, as illustrated byseal 169A ofFigure 4 . - Each
interface surface portion 181 can include an outwardly extendingretention feature 187. Theretention feature 187 is dimensioned to abut against surfaces of theengagement portion 179 to seat theouter air seal 169 during assembly and limit circumferential and/or radial movement. - The mounting
block 180 can be secured to the engine case 137 using one or more fasteners 163 (one shown inFigure 7 for illustrative purposes). Each mountingportion 182 defines anaperture 188 that receives a respective fastener 163 to mechanically attach the mountingportion 182 to the engine case 137 and limit relative movement of one or more outer air seals 169. In the illustrated example ofFigure 7 , the fastener 163 is a bolt, and theaperture 188 threadably receives a length of the bolt. In alternative examples, the fastener 163 is a clip or another structure to secure theouter air seal 169 to the engine static structure 136. - In the illustrated example of
Figure 7 , the adjacent outer air seals 169A, 169B are arranged in close proximity such that the respective mate faces 178A, 178B define an intersegment gap G that extends a distance in the circumferential direction T. In another example, referring toFigure 13 , mate faces 578A, 578B ofadjacent seals 569A, 569B are dimensioned to establish a ship lap joint to substantially reduce or eliminate the intersegment gap. Now referencing back toFigure 7 , the mountingblock 180 is situated between the mate faces 178A, 178B such that the mountingblock 180 spans across the intersegment gap G. A portion of the fastener 163 can be circumferentially aligned with one or more of the adjacent mate faces 178A, 178B and/or the intersegment gap G. Theinterface surface portions 181 abut against the adjacent outer air seals 169A, 169B to support the adjacent outer air seals 169A, 169B relative to the engine case 137 and limit circumferential and/or radial movement of the adjacent outer air seals 169A, 169B relative to the engine axis A. - The mounting
block 180 is arranged between theengagement portions block 180 secures theengagement portions block 180 is positioned at another location than along the intersegment gap G to secure theouter air seal 169 to the engine case 137, such as a midspan of theseal 169. - Referring back to
Figures 3-4 , with continued reference toFigures 5-9 , the outerair seal assembly 176 defines acooling arrangement 192 to deliver cooling flow F to portions of theouter air seal 169 and/or other portions of the outerair seal assembly 176 or components of the engine. The engine case 137 or another portion of the engine static structure 136 defines at least one or a plurality of coolingpassage apertures 139. Plenum(s) 174 extend between theimpingement face 189 and the engine case 137. Thecooling passage apertures 139 are configured to communicate with and receive cooling flow from cooling source(s) 175. - Each of the
cooling passage apertures 139 defines a respective passage axis PA that is oriented such that a projection of the passage axis PA intersects themain body 170 of theouter air seal 169. Thecooling passage apertures 139 eject cooling flow F into the plenum(s) 174 and toward themain body 170 of theouter air seal 169 in a direction along the passage axis PA. - The
main body 170 of theouter air seal 169 defines one or moreinlet flow apertures 191 to theinternal cavity 184. Theinlet flow apertures 191 deliver or otherwise communicate cooling flow F from thecooling source 175 to theinternal cavity 184 to cool adjacent portions of theseal 169 during engine operation. In the illustrated example ofFigures 3-7 , theimpingement face 189 defines inlets of theinlet flow apertures 191 to serve as conduits to provide cooling flow F into theinternal cavity 184. The inlets of theinlet flow apertures 191 may be distributed uniformly and/or non-uniformly in both and/or either the axial and/or circumferential directions X, T, for example. Theinlet flow apertures 191 may vary in cross sectional area in either the axial and/or circumferential streamwise directions in order to optimize the distribution of cooling flow F, within theinternal cavity 184, to achieve desired internal cavity Mach Number, Reynolds Number, and convective heat transfer, in order to tailor heat pickup and thermal cooling effectiveness requirements. Additionally the orientation and shape of theinlet flow apertures 191 may also comprise of both orthogonal and non-orthogonal apertures. In such instances theinlet flow apertures 191 may take the form of alternative geometries that may improve both thermal cooling and structural characteristics of theouter air seal 169.Inlet flow apertures 191 may consist of alternate geometric shapes and sizes, which may be, but are not limited to, conical, elliptical, and/or trapezoidal shapes of constant and/or varying cross sectional flow area and/or area ratio. - One or more of the
inlet flow apertures 191 can be spaced apart from each end wall of theengagement portion 179 that bounds theinternal cavity 184. Each of theinlet flow apertures 191 defines an aperture axis AA (Figure 3 ) such that a projection of the aperture axis AA intersects a wall of theinternal cavity 184. Theinlet flow apertures 191 are spaced from outlets of thecooling passage apertures 139. Theinlet flow apertures 191 are substantially exposed and free of any obstructions when in an assembled position such that theinlet flow apertures 191 communicate cooling flow F to theinternal cavity 184 during engine operation. Cooling flow F communicated by theinlet flow apertures 191 to theinternal cavity 184 can reduce relatively sharp and/or large thermal gradients resulting in locally high thermal strains that may otherwise occur in themain body 170 due to a lack of active flow or cooling to theinternal cavity 184. Reducing the potentially large radial and/or circumferential thermal gradients can mitigate detrimental thermal mechanical fatigue cracking resulting from high thermal-mechanical strains, which can improve the durability of outer air seals 169 comprising a ceramic or CMC material, for example. - The
inlet flow apertures 191 can be dimensioned to meter cooling flow F from thecooling source 175 into theinternal cavity 184 at a desired rate and reduce inefficiencies due to excess cooling flow F that may otherwise be communicated from thecooling source 175 to provide localized convective cooling in addition to a more optimal distribution of pressures withininternal cavity 184. The ability to effectively segregate, redirect, and redistribute internal flow and pressure withininternal cavity 184 can improve the utilization of available cooling flow F to achieve leakage and/or film cooling outflow requirements that may be needed to prevent ingestion and/or entrainment of the hot gases of the core flow path C. The tailoring and redistribution of cooling flow F can provide for a more efficientouter air seal 169 thermal cooling design configuration. By effectively utilizing less cooling flow F, improvements in thermodynamic cycle performance andouter air seal 169 durability capability can be achieved. -
Inlet flow apertures 191 may be used to locally tailor the distribution of cooling flow F to both axial and circumferntial portions ofinternal cavity 184. Additionally, the local internal pressure withininternal cavity 184 may be controlled and tailored to provide desired pressure ratio requirements tooutlet flow apertures Figure 8 ) and/or exposed mate faces 178 at either end of the outer air seal 169 (shown inFigure 6 ) in order to prevent ingestion and/or entrainment of the hot gases in the core flow path C. It may be desirable to keep the cooling flow F active and pressurized withincavity 184 in order to prevent contaminants and other environmental particulate from depositing itself in an otherwise stagnate cavity which may adversely impact the thermal mechanical characteristics of the material properties of theouter air seal 169. Unique tailoring of theinlet flow apertures 191 also can provide additional mitigation of local hot spots and sharp thermal gradients induced by variations in both local circumferential and axial heat flux and temperature distributions along the ID sealingsurface portion 177. - The
inlet flow apertures 191 can have a generally elliptical cross sectional profile and extend in a radial and/or predominately radial direction R through a thickness of a wall of themain body 170. Additionally,inlet flow apertures 191 may be non-orthogonal and oriented in both a circumferential and/or axial direction. Alternativeinlet flow apertures 191 geometries may be utilized either independently and/or in conjunction with, including, but not limited to, slots, conical, diffused, single or multi-lobed, shapes and/or consist of more complex geometric shapes comprising of single or multiple concave and/or convex surfaces. - Referring to
Figures 3 and 4 , theinlet flow apertures 191 can be arranged relative to thecooling passage apertures 139 to provide localized cooling to selected portions of theouter air seal 169. The projection of the passage axis PA of at least somecooling passage apertures 139 can intersect theimpingement face 189 to provide localized impingement cooling, but the projection of the passage axis PA of othercooling passage apertures 139 may not. Each passage axis PA can be perpendicular or transverse to theimpingement face 189 at the point of intersection (see, e.g., passage apertures 139A and 139B ofFigure 3 , respectively). The passage axis PA of at least some of thecooling passage apertures 139 can be aligned with or otherwise intersect a respective one of the inlets 191 (see, e.g., passage apertures 139A) to increase an impingement distance between an outlet of therespective passage apertures 139 and theseal 169. The increased impingement distance can reduce the heat transfer augmentation, which can reduce thermal gradients and more uniformly distribute the cooling flow F in the radial direction R. - Some of the
inlet flow apertures 191 can be offset from a projection of the passage axis PA of thecooling passage apertures 139, as illustrated by inlets 191' inFigure 3 (shown in dashed lines). Offsetting the inlet flow apertures 191' allows the cooling flow F to follow along surfaces of theimpingement face 189 to cool adjacent portions of theouter air seal 169 prior to entry into the inlet flow apertures 191'. Aligning some of thecooling passage apertures 139 with respectiveinlet flow apertures 191 but offsetting othercooling passage apertures 139 with respect to each of theinlet flow apertures 191 can provide a mixture of localized, impingement cooling to surfaces of theimpingement face 189 and more direct cooling to theinternal cavity 184. In examples, eachinlet flow aperture 191 is aligned with a respectivecooling passage apertures 139. In other examples, eachinlet flow aperture 191 is offset from eachcooling passage apertures 139. - The
outer air seal 169 is exposed to temperature differentials between the gases in the core flow path C and the cooling flow F from thecooling source 175. Portions of theouter air seal 169 may be exposed to different thermal loads, which can result in thermal gradients across theouter air seal 169. The thermal gradients may cause cracks to form in theouter air seal 169, for example. - Referring to
Figure 8 , thebaffle 190 is secured to theseal 169 in at least a portion of theinternal cavity 184.Figure 3 illustrates theseal assembly 176 with thebaffle 190 uninstalled in theinternal cavity 184. Thebaffle 190 can be dimensioned to be removably inserted into theinternal cavity 184. In other examples, thebaffle 190 is permanently secured to in theinternal cavity 184, such as by molding thebaffle 190 in themain body 170, for example. - The
baffle 190 divides theinternal cavity 184 into a plurality of localized regions. In the illustrated example ofFigure 8 , thebaffle 190 divides theinternal cavity 184 into at least afirst region 195 and asecond region 196 that is separate and distinct from thefirst region 195. Thefirst region 195 serves as a cooling channel to direct or otherwise guide the cooling flow F along a perimeter of theinternal cavity 184. Thefirst region 195 can provide an increase in the max flux as well as optimized convective heat transfer coefficient (HTC) augmentation to radially inner portions of theinternal cavity 184 adjacent to theseal face surface 183. Thefirst region 195 can be established along portions of theouter air seal 169 that may be more susceptible to thermal loads or gradients, thereby establishing a prioritized cooling scheme. The prioritized cooling can reduce thermal gradients in theouter air seal 169, thereby reducing a likelihood of cracking and improving durability. - The
regions outer air seal 169 with respect to expected heat loads and thermal gradients that may be experienced during engine operation. In some examples, thefirst region 195 defines a first volume, thesecond region 196 defines a second volume, and the first volume is less than half of the second volume. In further examples, the first volume is less than 10% or 25% of the second volume. - At least one
inlet flow apertures 191 is defined along thefirst region 195. Theinlet flow apertures 191 can be defined along theimpingement face 189, for example. Thefirst region 195 communicates cooling flow F from theinlet flow apertures 191 to one or moreoutlet flow apertures 197 defined by themain body 170. In the illustrated example ofFigure 8 , one or more outlet flow apertures 197 (one shown for illustrative purposes) are defined along the trailingedge portion 194 of themain body 170. In some examples, the sealingsurface portion 177 defines one or more outlet flow apertures 197' (shown in dashed lines) that serve as film cooling holes to eject cooling flow F from theinternal cavity 184 into the core flow path C to cool adjacent surfaces of the outer airseal face surface 183. In other examples, one or moreoutlet flow apertures 197" (one shown for illustrative purposes) are defined along theleading edge portion 193 of theseal 169. Theoutlet flow apertures 197" can reduce thermal gradients across a thickness of the walls of theleading edge portion 193. - The
baffle 190 is arranged in theinternal cavity 184 to change a direction of the cooling flow F from theinlet flow aperture 191 relative to the respective aperture axis AA. In the illustrated example ofFigures 8 and9 , thebaffle 190 includes an elongated baffle body 190-1 having a generally C-shaped geometry. A first end 190-2 of thebaffle 190 is contoured and is dimensioned to abut against surfaces of theinternal cavity 184 along walls of themain body 170 that define theimpingement face 189. Thebaffle 190 fluidly isolates the first andsecond regions baffle 190 can be substantially planar. In some examples, second end 190-3' (shown in dashed lines inFigure 9 ) is contoured and is dimensioned to abut against surfaces of theinternal cavity 184 such that thebaffle 190 fluidly isolates the first andsecond regions - In the illustrated example of
Figures 8 and9 , thebaffle 190 is substantially free of any impingement holes such that thefirst region 195 is fluidly isolated from thesecond region 196 between theinlet flow apertures 191 and theoutlet flow apertures 197. For the purposes of this disclosure, the term "substantially free" means that less than 3% of the baffle volume defines holes. In other examples, thebaffle 190 includes one or more impingement holes 199 (shown in dashed lines inFigure 8 ). Although theoutlet flow apertures 199 described herein are referred to as impingement holes it is to be understood that these features may consist of noncircular and cylindrical shapes, such as but not limited to, slot, elliptical, oval, race track, or other multi-faceted geometry shapes comprising of conical and/or convex surfaces. - In the illustrated example of
Figure 7 , ends of the baffle 190 (shown in dashed lines) are spaced apart from walls of theinternal cavity 184 such that thefirst region 195 and/or thesecond region 196 includesoutlets 198 along the mate faces 178. Thebaffle 190 can include one or more bumpers or protrusions 190-4 (Figure 9 ) such as pedestals to space apart thebaffle 190 and offset it from the surfaces of theinternal cavity 184, creating a gap or channel by which cooling flow F may be regulated and/or controlled and then dispersed throughoutlet flow apertures Figure 8 . - The mounting
block 180 can be dimensioned relative to surfaces of theinternal cavity 184 such that cooling flow F in theinternal cavity 184 can exit from theoutlets 198, intoopenings 185 along the mate faces 178, and into the intersegment gaps G. Communication of the cooling flow F from theoutlets 198 can cool portions of the mate faces 178, thereby improving durability of the outer air seals 169. Communicating the cooling flow F into each intersegment gap G creates a fluidic sealing boundary and cooling purge flow relationship to be formed, which can reduce the likelihood of ingestion of hot combustion gases from the core flow path C and into the intersegment gap G. In other examples, ends of thebaffle 190 are dimensioned to abut against the walls of theinternal cavity 184 such that thefirst region 195 is substantially fluidly isolated from the mate faces 178. - The
baffle 190 can be dimensioned such that thefirst region 195 has a circuitous path between surfaces of thebaffle 190 and walls of themain body 170 that define theinternal cavity 184. Thebaffle 190 directs or guides the cooling flow F from theinlet flow apertures 191 along the circuitous path to provide localized cooling to adjacent portions of themain body 170. - In the illustrated example of
Figure 8 , thefirst region 195 includes a circuitous path that follows a perimeter of theinternal cavity 184. The circuitous path of the first region is defined by at least first, second and third sections 195-1, 195-2, 195-3. Thebaffle 190 can be arranged to establish a relatively small offset from walls of theinternal cavity 184 to improve heat transfer augmentation to portions of theseal 169 adjacent to the first, second and third sections 195-1, 195-2, 195-3. For example, an offset or width W1 of thebaffle 190 from surfaces of theinternal cavity 184 along thefirst region 195 can be less than 5% or 10% of a height H and/or length L of theinternal cavity 184. The relatively small offset can increase velocities of the cooling flow F along thefirst region 195. - The width W1 along the
first region 195 can vary for each of the first, second and third sections 195-1, 195-2, 195-3. The varying offset along thefirst region 195 can change the Mach No. distribution of the cooling flow along the flow path to provide prioritized cooling augmentation to surfaces of theseal 169 along thefirst region 195. For example, width W1-1 along the first section 195-1 can be greater than width W1-2 along the second section 195-2 to cause relatively greater cooling flow F to circulate along the first section 195-1. The relatively greater width W1-1 can reduce a velocity of the cooling flow F, thereby increasing the local static pressure within the channel between thebaffle 190 and the wall surfaces ofinternal cavity 184, along the first section 195-1, and the relatively lesser width W1-2 can the decrease a static pressure of the cooling flow F to reduce the pressure ratio across theoutlet flow apertures 197, 197' for improved film cooling. The gap between thebaffle 190 and the wall surfaces ofinternal cavity 184 may be set by altering the channel heights W1-1 and W1-2 in order to achieve the desired local cooling flow F velocity and required local static pressure to ensure sufficient driving pressure exists acrossoutlet flow apertures out air seal 169. - The first section 195-1 extends transversely from the second and third sections 195-2, 195-3 such that a cross section of the
first region 195 has a generally C-shaped geometry. The first section 195-1 has a component in the radial direction R with respect to the engine axis A, for example. The first section 195-1 extends along walls of theleading edge portion 193 that define theinternal cavity 184. Thebaffle 190 can be situated in other orientations. For example, an orientation of thebaffle 190 can be inverse to the orientation of thebaffle 190 illustrated inFigure 8 such that the first section 195-1 extends along walls of the trailingedge portion 194 that define theinternal cavity 184. - The second section 195-2 has a component in the axial direction X with respect to the engine axis A, for example, such that the second section 195-2 is defined between the
baffle 190 and walls of themain body 170 defining theseal face surface 183. - The third section 195-3 has a component in the axial direction X with respect to the engine axis A, for example, such that the third section 195-3 is defined between the
baffle 190 and walls of themain body 170 defining theimpingement face 189. The third section 195-3 fluidically connects the first section 195-1 and theinlets 191. The third section 195-3 extends along walls of themain body 170 that define theimpingement face 189 such that thefirst region 195 at least partially surrounds thesecond region 196. In the illustrative example ofFigure 8 , thesecond region 196 extends along sidewalls of the trailingedge portion 194, but is spaced apart from walls of theleading edge portion 193. A length of the third section 195-3 can be less than, equal to, or greater than a length of the second section 195-2 with respect to the axial direction X or engine axis A depending on flow, heat transfer and pressure loss requirements needed to minimize local thermal gradients while ensure positive pressure ratio and outflow conditions foroutlet flow apertures Figure 8 . The location and number of flow apertures can be selected depending onouter air seal 169 cooling flow F allocations and durability life requirements. - The first section 195-1 fluidically connects the second section 195-2 and the
inlet flow apertures 191 to deliver cooling flow F from thecooling source 175 to the second section 195-2. The second section 195-2 fluidically connects the first section 195-1 and one ormore outlets 197 defined by themain body 170 to deliver the cooling flow F to downstream portions of theouter air seal 169. In some examples, the cooling flow F is ejected fromoutlets 198 of thefirst region 195 along the mate faces 178, as illustrated byFigure 7 . In the illustrated example ofFigure 8 , the second end 190-3 of thebaffle 190 is spaced apart from walls of theinternal cavity 184 such that cooling flow F is communicated from the second section 195-2 to thesecond region 196 subsequent to be guided along the circuitous path established by thefirst region 195. - Various materials can be utilized to manufacture the
outer air seal 169, mountingblock 180 andbaffle 190. In some examples, theouter air seal 169 is made of a first material, and the mountingblock 180 and/or baffle 190 is made of a second, different material. For example, the first material can include a ceramic or ceramic matrix composite (CMC) material. Theouter air seal 169 can be formed from one or more layers L of a CMC layup (Figure 3 ). Theouter air seal 169 can be made of another material, such as a high temperature metal, alloy, or composite material. The mountingblock 180 and/or baffle 190 can be made of a second material such as a high temperature composite, metal, or alloy, such as a nickel or cobalt-based superalloy, for example. Surfaces of thebaffle 190 can include one or more coatings, such as to reduce interaction between the CMC material of theouter air seal 169 and the nickel alloy material of thebaffle 190. The first and second materials can differ. In other examples, theouter air seal 169 is made of a first material, and the mountingblock 180 and/or baffle 190 is made of a second material that is the same as the first material, including any of the materials disclosed herein. Theouter air seal 169 is formed to have a unitary construction. In alternative examples, the sealingsurface portion 177 and eachengagement portion 179 are separate and distinct components that are mechanically attached to one another with one or more fasteners. In some examples, thebaffle 190 is made from sheet metal, or other materials with moderate temperature capability, which may include but not limited to Inconel, Cobalt-Chrome, Hastelloy, Nickel etc. and is formed to the desired geometry shape. - The inlet and
outlet flow apertures main body 170. Alternatively, the inlet andoutlet apertures main body 170 by arranging the layers L of the CMC layup (Figure 3 ), for example. - A method of sealing is as follows. With reference to
Figures 7 and8 , thebaffle 190 is inserted into anopening 185 to theinternal cavity 184 such that thebaffle 190 is secured to theouter air seal 169. A mountingblock 180 is positioned between mate faces 178A, 178B of each adjacent pair ofseals interface surface portions engagement portions mounting block 180 to secure the outerair seal assembly 176 to the engine static structure 136. - Referring to
Figures 3-4 and8 , with continued reference toFigure 7 , during operation of the engine, pressurized cooling flow F is communicated from thecooling source 175 to thecooling passage apertures 139. Thecooling passages apertures 139 eject the cooling flow F into theplenum 174 and in a direction toward theimpingement face 189 of theseal 169. The cooling flow F can be discharged from thecooling passage apertures 139 at a predetermined pressure and velocity such that the cooling flow F impinges on localized surfaces of theimpingement face 189. The cooling flow F circulates from theplenum 174 and into theinlet apertures 191. The cooling flow F circulates from theinlet flow apertures 191 and disperses into the third section 195-3 of thefirst region 195. The cooling flow F can circulate from the inlet aperture(s) 191 to the third section 195-3, then to the first section 195-1, and then to the second section 195-2. The cooling flow F can be ejected from one ormore outlets 197, 197' along the first and/or second sections 195-1, 195-2 and into the gas path such as the core flow path C. In examples, the cooling flow F circulates from theplenum 174 into the inlet flow aperture(s) 191', and then into thesecond region 196. -
Figure 10 illustrates abaffle 290 according to another example. Thebaffle 290 can have a non-uniform offset from surfaces of theinternal cavity 284. In the illustrated example ofFigure 10 , walls of baffle body 290-1 are contoured such that a distance or width offirst region 295 between thebaffle 290 and walls ofmain body 270 varies at locations along thefirst region 295 in a first direction D1, which can be at least one of the circumferential, radial and/or axial directions T, R, X, and/or a combination thereof (seeFigures 3-4 and8 ), and wherein a passage axis of the inlet cooling aperture(s) (see, e.g., passage axis PA ofFigure 8 ) can be oriented orthogonally or non-orthogonally with respect to grooves 290-4 to minimize or otherwise reduce pressure losses, for example. - The baffle body 290-1 defines one or more grooves 290-4 such that an offset or width W1 between surfaces of the grooves 290-4 and the
main body 270 is greater than a width W2 between themain body 270 and other portions of the baffle body 290-1 along thefirst region 295. A cross section of the grooves 290-4 can have a substantially rectangular profile, as illustrated byFigure 10 . The grooves can have other geometries, such as a generally concave geometry as illustrated by grooves 390-4 ofFigure 11 . The features ofFigures 10 and 11 can be incorporated into thebaffle 190 along the first, second and/or third sections 195-1, 195-2, 195-3 of thefirst region 195 ofFigure 8 , for example. The grooves 290-4/390-4 can be circumferentially aligned with respect to injection nozzles of thecombustor 56 orvanes 70 axially forward of theouter air seal 169, which may cause non-uniform thermal loads on theouter air seal 169, for example. -
Figure 12 illustrates adjacent seal assemblies 476 according to another example. Two adjacent seals are labeled as 469A, 469B. Eachbaffle 490 is secured between arespective seal 469A/469B and surfaces of mountingblock 480. In the illustrated example ofFigures 12 ,baffle 490 is dimensioned such that portions of a baffle body 490-1 of thebaffle 490, such as a portion of first end 490-2, are trapped between interface portions 481 of the mountingblock 480 and ramped surfaces 486 of internal cavity 484 defined by the respective seal 469 that are opposed to surfaces of the respective interface portion 481 to limit relative movement in the axial, radial and/or circumferential directions X, R, T, for example. - It should be understood that relative positional terms such as "forward," "aft," "upper," "lower," "above," "below," and the like are with reference to the normal operational attitude of the vehicle and should not be considered otherwise limiting.
- Although the different examples have the specific components shown in the illustrations, embodiments of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from one of the examples in combination with features or components from another one of the examples.
- Although particular step sequences are shown, described, and claimed, it should be understood that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present disclosure.
- The foregoing description is exemplary rather than defined by the limitations within. Various non-limiting embodiments are disclosed herein, however, one of ordinary skill in the art would recognize that various modifications and variations in light of the above teachings will fall within the scope of the appended claims. It is therefore to be understood that within the scope of the appended claims, the disclosure may be practiced other than as specifically described. For that reason the appended claims should be studied to determine true scope and content.
Claims (15)
- A seal assembly (76,176) for a gas turbine engine (20) comprising:a blade outer air seal (BOAS) (169) including an elongated seal body (170) having a seal face (183) for bounding a gas path and an opposed impingement face (189), the seal body (170) defining an internal cavity (184) extending in a circumferential direction (T) between opposed mate faces (178) and extending in a radial direction (R) between walls of the seal body (170) defining the seal face (183) and impingement face (189); anda baffle (190;290) that divides the internal cavity (184) into at least a first region (195;295) and a second region (196;296), the first region (195;295) including a first section (195-1) extending transversely from a second section (195-2), the first section (195-1) having a component in the radial direction (R), and the second section (195-2) having a component in an axial direction (X) such that the second section (195-2) is defined between the baffle (190;290) and the walls of the seal body (170) defining the seal face (183).
- The seal assembly (76,176) as recited in claim 1, wherein the first region (195;295) defines a first volume, the second region (196;296) defines a second volume, and the first volume is less than half of the second volume.
- The seal assembly (76,176) as recited in claim 1 or 2, wherein the first section (195-1) interconnects the second section (195-2) and one or more inlet flow apertures (191) defined by the seal body (170).
- The seal assembly (76,176) as recited in claim 3, wherein the second section (195-2) interconnects the first section (195-1) and one or more outlet flow apertures (197,197',197") defined by the seal body (170).
- The seal assembly (76,176) of claim 4, wherein the baffle (190;290) is substantially free of impingement holes (199) such that the first region (195;295) is fluidly isolated from the second region (196;296) between the one or more inlet flow apertures (191) and the one or more outlet flow apertures (197,197',197").
- The seal assembly (76,176) as recited in claim 3, 4 or 5, wherein the first region (195;295) includes a third section (195-3) that extends transversely from the first section (195-1) such that the third section (195-3) has a component in the axial direction (X) to interconnect the first section (195-1) and the one or more inlet flow apertures (191), optionally wherein the third section (195-3) extends along walls of the seal body (170) that define the impingement face (189) such that the first region (195;295) at least partially surrounds the second region (196;296).
- The seal assembly (76,176) as recited in any preceding claim, wherein an offset of the baffle (190;290) from surfaces of the internal cavity (184) along the first region (195) varies at locations along the baffle (190;290) in at least one of the circumferential direction (T) and the axial direction (X).
- The seal assembly (76,176) as recited in any preceding claim, wherein the first section (195-1) extends along walls of a leading edge region (193) of the seal body (170) that defines the internal cavity (184).
- The seal assembly (76,176) as recited in any preceding claim, further comprising a pair of mounting blocks (180;480) insertable into respective openings along the mate faces (178) to secure the blade outer air seal (169) to an engine static structure (136).
- The seal assembly (76,176) as recited in claim 9, wherein the internal cavity (184) extends circumferentially from one of the openings to another one of the openings.
- The seal assembly (76,176) as recited in any preceding claim, wherein the first region (195;295) includes outlet flow apertures (198) along the mate faces (178).
- The seal assembly (76,176) as recited in any preceding claim, wherein the blade outer air seal (169) is made of a first material including a ceramic material, and the seal has a unitary construction.
- A method of sealing of a gas turbine engine (20), comprising:securing a seal assembly (76; 176) to an engine static structure (136), wherein the seal assembly (76; 176) includes a blade outer air seal (169) and a baffle (190;290), the blade outer air seal (169) includes an elongated seal body (170) having a seal face (183) that bounds a gas path and an opposed impingement face (189), the seal body (170) defines an internal cavity (184) that extends in a circumferential direction (T) between opposed mate faces (178) and extends in a radial direction (R) between walls of the seal body (170) defining the seal face (183) and impingement face (189), the baffle (190;290) divides the internal cavity (184) into at least a first region (195; 295) and a second region (196;296), the first region (195;295) includes a first section (195-1) extending transversely from a second section (195-2), the first section (195-1) has a component in a radial direction (R), and the second section (195-2) has a component in an axial direction (X) such that the second section (195-2) is defined between the baffle (190;290) and walls of the seal body (170) defining the seal face (183);communicating cooling flow from at least one inlet flow aperture (191) to the first section (195-1), and then to the second section (195-2), the at least one inlet flow aperture (191) defined along the impingement face (189); andejecting the cooling flow from the second section (195-2) into the gas path.
- The method as recited in claim 13, wherein the first region (195;295) includes a third section (195-3) that extends transversely from the first section (195-1) such that the third section (195-3) has a component in the axial direction (X) to interconnect the first section (195-1) and the at least one inlet flow aperture (191), the third section (195-3) extending along walls of the seal body (170) that define the impingement face (189) such that the first region (195;295) at least partially surrounds the second region (196;296), optionally an end of the baffle (190;290) abuts against the walls of the seal body (170) defining the impingement face (189) and/or the engine static structure (136) is an engine case (137) that defines a plurality of cooling passages (139), and each of the plurality of cooling passages (139) defines a passage axis (PA) that is oriented such that a projection of the passage axis (PA) intersects the impingement face (189).
- The method as recited in claim 13 or 14, wherein the step of ejecting the cooling flow includes ejecting the cooling flow from outlet flow apertures (197,197',197") of the first region (195;295) along the mate faces (178).
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US15/982,062 US11242764B2 (en) | 2018-05-17 | 2018-05-17 | Seal assembly with baffle for gas turbine engine |
Publications (3)
Publication Number | Publication Date |
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EP3569823A2 EP3569823A2 (en) | 2019-11-20 |
EP3569823A3 EP3569823A3 (en) | 2020-01-15 |
EP3569823B1 true EP3569823B1 (en) | 2021-06-30 |
Family
ID=66589445
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
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EP19175130.4A Active EP3569823B1 (en) | 2018-05-17 | 2019-05-17 | Seal assembly with baffle for gas turbine engine |
Country Status (2)
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US (1) | US11242764B2 (en) |
EP (1) | EP3569823B1 (en) |
Families Citing this family (2)
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US10634010B2 (en) * | 2018-09-05 | 2020-04-28 | United Technologies Corporation | CMC BOAS axial retaining clip |
US10648407B2 (en) * | 2018-09-05 | 2020-05-12 | United Technologies Corporation | CMC boas cooling air flow guide |
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FR2416345A1 (en) * | 1978-01-31 | 1979-08-31 | Snecma | IMPACT COOLING DEVICE FOR TURBINE SEGMENTS OF A TURBOREACTOR |
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US5165847A (en) * | 1991-05-20 | 1992-11-24 | General Electric Company | Tapered enlargement metering inlet channel for a shroud cooling assembly of gas turbine engines |
US7740444B2 (en) * | 2006-11-30 | 2010-06-22 | General Electric Company | Methods and system for cooling integral turbine shround assemblies |
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US7665953B2 (en) * | 2006-11-30 | 2010-02-23 | General Electric Company | Methods and system for recuperated cooling of integral turbine nozzle and shroud assemblies |
GB0703827D0 (en) * | 2007-02-28 | 2007-04-11 | Rolls Royce Plc | Rotor seal segment |
FR2922589B1 (en) * | 2007-10-22 | 2009-12-04 | Snecma | CONTROL OF THE AUBES SET IN A HIGH-PRESSURE TURBINE TURBINE |
FR2925109B1 (en) * | 2007-12-14 | 2015-05-15 | Snecma | TURBOMACHINE MODULE PROVIDED WITH A DEVICE FOR IMPROVING RADIAL GAMES |
GB201112163D0 (en) * | 2011-07-15 | 2011-08-31 | Rolls Royce Plc | Tip clearance control for turbine blades |
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US9500095B2 (en) * | 2013-03-13 | 2016-11-22 | Pratt & Whitney Canada Corp. | Turbine shroud segment sealing |
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CA2916710A1 (en) * | 2015-01-29 | 2016-07-29 | Rolls-Royce Corporation | Seals for gas turbine engines |
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US10968761B2 (en) * | 2018-11-08 | 2021-04-06 | Raytheon Technologies Corporation | Seal assembly with impingement seal plate |
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US11352897B2 (en) * | 2019-09-26 | 2022-06-07 | Raytheon Technologies Corporation | Double box composite seal assembly for gas turbine engine |
US11220924B2 (en) * | 2019-09-26 | 2022-01-11 | Raytheon Technologies Corporation | Double box composite seal assembly with insert for gas turbine engine |
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- 2018-05-17 US US15/982,062 patent/US11242764B2/en active Active
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Non-Patent Citations (1)
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US20190353045A1 (en) | 2019-11-21 |
US11242764B2 (en) | 2022-02-08 |
EP3569823A3 (en) | 2020-01-15 |
EP3569823A2 (en) | 2019-11-20 |
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